Gas turbine engines may be used to power various types of vehicles and systems, such as air or land-based vehicles. In typical gas turbine engines, compressed air generated by axial and/or radial compressors is mixed with fuel and burned, and the expanding hot combustion gases are directed along a flowpath and through a turbine nozzle having stationary turbine vanes. The gas flow deflects off of the vanes and impinges upon blades of a turbine rotor. A rotatable turbine disk or wheel, from which the turbine blades extend, spins at high speeds to produce power. Gas turbine engines used in aircraft use the gas turbine aft end to produce a forward thrust. Other gas turbine engines may use the power to turn a propeller or an electrical generator. The gas turbine engine component comprising a wall with a wall surface configured to be exposed to the hot combustion gas flow during engine operation (i.e., a “gas path surface”) and an opposing wall surface is referred to herein as a “gas path component.”
One way to increase cycle efficiency, power density, and fuel efficiency of a gas turbine engine is to operate at higher turbine inlet temperatures (TIT). In most engines, the turbine inlet temperatures have increased well above the metallurgical limit of gas path components, making the gas path components (e.g., turbine nozzles and turbine blades and vanes) more susceptible to oxidation and thermo-mechanical fatigue. Film and impingement cooling of conventional gas path components are widely used techniques that allow higher turbine inlet temperatures by maintaining material temperatures within acceptable limits. For example, with film cooling, air may be extracted from the compressor and forced through internal cooling passages within the gas path component (e.g., a turbine blade) before being ejected through discrete film cooling holes onto an outer wall surface thereof (in this example, the gas path surface). The cooling air leaving these film cooling holes forms a film layer of cooling air on the outer wall surface, the film layer protecting the component from hot gas exiting the combustor by substantially reducing heat transfer from the hot gas to the gas path surface as the cooling air is at a lower temperature than the hot gas. Turbulence promoters and pin fins in a cooling medium flowpath through the component may also or alternatively be used to improve cooling. Although the aforementioned cooling techniques operate adequately, they may be improved for higher cooling effectiveness. Unfortunately, achieving a high heat transfer rate often involves large temperature gradients (that decrease component life) and complex geometries that are difficult and expensive to manufacture, using complex, fragile, and convectively augmented ceramic cores.
Accordingly, it is desirable to provide gas path components of gas turbine engines, such as turbine blade airfoils, and methods for cooling the same. Furthermore, other desirable features and characteristics of the inventive subject matter will become apparent from the subsequent detailed description of the inventive subject matter and the appended claims, taken in conjunction with the accompanying drawings and this background of the inventive subject matter.